ExoCube is the latest National Science Foundation (NSF) funded space weather CubeSat and is a collaboration between PolySat, Scientific Solutions Inc. (SSI), the University of Wisconsin, NASA Goddard and SRI International. The 3U will carry a mass spectrometer sensor suite, EXOS, in to low earth orbit (LEO) to measure neutral and ionized particles in the exosphere and thermosphere. Measurements of neutral and ion particles are directly impacted by the angle at which they enter EXOS and which leads to pointing requirements. A combination of a gravity gradient system with a momentum bias wheel is proposed to meet pointing requirements while reducing power requirements and overall system complexity. A MATLAB simulation of dynamic and kinematic behavior of the system in orbit is implemented to guide system design and verify that the pointing requirements will be met. The problem of achieving the required three-axis pointing is broken into four phases: detumbling, initial attitude acquisition, wheel spin-up, and attitude maintenance. Ultimately, this configuration for attitude control in a CubeSat could be applied to many future missions with the simulation serving as a design tool for CubeSat developers.

This thesis examines the feasibility of substituting the system input script of Cal Poly’s Horizon Simulation Framework (HSF) with a Model Based Systems Engineering (MBSE) model designed with the Systems Modeling Language (SysML). A concurrent student project, SysML Output Interface Creation for the Horizon Simulation Framework, focused on design of the HSF Translator Plugin which converts SysML models to an HSF specific XML format. A SysML model of the HSF test case, Aeolus, was designed. The original Aeolus HSF input script and the translated SysML input script retained the format and dependency structure required by HSF. Both input scripts returned identical results and thus validated the feasibility of linking SysML with HSF through the HSF Translator Plugin. A second SysML model of the Cal Poly CubeSat mission, ExoCube, was also designed and converted into an HSF input script. The ExoCube input script also retained the format and dependency structure required by HSF. This demonstrated that future SysML models can be used in conjunction with the HSF Translator Plugin to create a functional HSF system input script.

There is an increased need for spacecraft to quickly and efficiently de-orbit themselves as the amount of debris in orbit around Earth grows. Defunct spacecraft pose a significant threat to the LEO environment due to their risk of fragmentation. If these spacecraft are de-orbited at the end of their useful life their risk to future spacecraft is greatly lessened. A proposed method of efficiently de-orbiting spacecraft is to use an inflatable thin-film envelope to increase the body's area to mass ratio and thusly shortening its orbital lifetime. The system and analysis presented in this project is sized for use on a CubeSat as they are an effective utility as a technology demonstration platform. Analysis has been performed to characterize the orbital dynamics of high area to mass ratio spacecraft as well as the leak rate of such an inflatable device in a vacuum environment. Results show that a 1U CubeSat can be de-orbited using a 1.7 meter diameter spherical device in just under one year while using 0.7 grams of inflating gas, this is compared to over 25 years without any method of post-mission disposal.

Spacecraft operating on the Martian surface have used relay satellites as a means of improving communication capabilities, mainly in terms of bandwidth and availability. However, the spacecraft used to achieve this have been large spacecraft (1000s of kilograms) and were not designed with relay capability as the design priority. This thesis explores the possibility of using a CubeSat-based constellation as a communications network for spacecraft operating on the Martian surface. Brute-force techniques are employed to explore the design space of possible constellations. An analysis of constellation configurations that provide complete, continuous coverage of the Martian surface is presented. The stability of these constellations are analyzed, and recommendations are made for stable configurations and the orbital maintenance thereof. Link budget analysis is used to determine the communications capability of each constellation, and recommendations are made for sizing each communication element. The results of these three analyses are synthesized to create an architecture generation tool. This tool is used to identify mission architectures that suit a variety of mission requirements, and these architectures are presented. The primary recommended architecture utilizes 18 CubeSats in three orbital planes with six additional larger relay satellites to provide an average of over one terabit/sol downlink and 100 kbps uplink capability.

Attitude determination and estimation algorithms are developed and implemented in simulation for the Exocube satellite currently under development by PolySat at Cal Poly. A mission requirement of ±5˚ of attitude knowledge has been flowed down from the NASA Goddard developed payload, and this requirement is to be met with a basic sensor suite and the appropriate algorithms. The algorithms selected in this work are TRIAD and an Extended Kalman Filter, both of which are placed in a simulation structure along with models for orbit propagation, spacecraft kinematics and dynamics, and sensor and reference vector models. Errors inherent from sensors, orbit position knowledge, and reference vector generation are modeled as well. Simulations are then run for anticipated dynamic states of Exocube while varying parameters for the spacecraft, attitude algorithms, and level of error. The nominal case shows steady state convergence to within 1˚ of attitude knowledge, with sensor errors set to 3.5˚ and reference vector errors set to 2˚. The algorithms employed have their functionality confirmed with the use of STK, and the simulations have been structured to be used as tools to help evaluate attitude knowledge capabilities for the Exocube mission and future PolySat missions.

This thesis is the product of an effort to develop a CubeSat Conceptual Design Tool for the California Polytechnic State University CubeSat Laboratory. Such a tool is necessary due to inefficiencies with the current conceptual design process. It is being developed to increase accessibility, reduce design time, and promote good systems engineering within CubeSat development. The development of the architecture of a conceptual design tool, the core user-interface element, and the completion of a module for the electrical power subsystem is the focus of this thesis. The architecture is built around different modules to design different subsystems that work in conjunction. The module in the tool was developed to allow a user to size an electrical power subsystem, and that is the basis for future subsystem development. Model-based Systems Engineering was also utilized as an endpoint for the tool’s outputs, and a CubeSat Model has been built for this effort. Validation has been successful on the Conceptual Design Tool as implemented at this time, so the tool it is ready to design CubeSat electrical power subsystems and be expanded upon by other tool developers.

The Miniaturized Ultraviolet Imager (MUVI), is a compact wide field UV imaging instrument in development at UC Berkeley Space Sciences Laboratory and Cal Poly, San Luis Obispo. MUVI is designed to fit in a 2U CubeSat form factor and provide wide field, high resolution images of the ionosphere at far ultraviolet wavelengths. This thesis details the design and analyses of MUVI’s deployable cover mirror mounting flexures. Three different flexure geometries were evaluated, an optimal candidate was determined based on a number of criteria including isolation of vibration and stress to the mirrors, manufacturability and cost. The design of the flexure system includes the flexure blades themselves, Invar pads bonded to the mirror to mitigate the difference in CTEs of the different material, mounting of flexure blades to the deployable cover and ground support equipment for assembly and testing. During the design of the flexures, various materials were studied, and Titanium was concluded as the optimum material due to its combination of high strength and flexibility compared to stainless steel, aluminum and other metals. Utilizing titanium, several flexure designs were proposed, and three candidates were selected to be manufactured and tested. Throughout the design phase, all flexures went through several rounds of analysis utilizing finite element analysis to simulate quasi-static loads, modal analysis of the systems natural frequency as well as random vibration simulations to simulate testing environments. Once the front-runner designs were selected and manufactured, several tests were conducted. Testing included adhesive bond coupon testing of the adhesive in tension and bending to experimentally validate the bonding size of the invar pads would be sufficient. The adhesive bond testing conducted tension and three-point bend tests to characterize the epoxy adhesive used in the flexure assembly. Testing also consisted of sine sweep and random vibration environment in accordance with the NASA General Environmental Verification Standard to qualify the hardware for spaceflight. Throughout the vibration testing, an autocollimator was used pre and post-test to measure shifts in the optical alignment of the mirror after it underwent vibration qualification testing. Experimental and analytical models were compared once all testing was completed. The Curved Blade showed to test in the real world very close to that predicted by the finite element model, however, the Bent Blade and Z Blade showed a larger difference between analysis and test. Discussion into the reasoning for this difference and lessons learned is included.

The rising complexity of CubeSat missions and the unique challenges faced by educational CubeSat programs lead to high rates of mission failure. Implementing project management and systems engineering practices can alleviate these challenges and improve mission success rates for educational CubeSat developers. However, existing project management and systems engineering resources are too cumbersome and often assume the student has a base-level understanding of project management and systems engineering fundamentals. A new universal project management and systems engineering framework was created and tailored specifically to the needs of an educational CubeSat mission. The framework was designed to accommodate first-time CubeSat developers, has no base-level assumptions, and uses software accessible by the majority of university CubeSat programs. The framework was implemented on an educational CubeSat mission being designed by a first-time CubeSat developer. The framework was iteratively updated based on the developer’s feedback and experience using the templates, tools, and trainings. The universal project management and systems engineering framework for educational CubeSat missions demonstrated that it was effectively tailored to the needs of educational CubeSat missions, taught students the value of project management and systems engineering, enhanced students’ professional development, and was accessible for first-year undergraduate students to utilize with minimal intervention.

A small satellite can be adversely affected by Earth’s magnetic field due to the resulting torque the magnetic field exerts on the satellites magnetic dipole moment. Therefore, this dipole moment needs to be estimated during the development of the satellite to make sure that the torque does not become a problem once the satellite is in orbit. This needs to be done for the MIniture STudent satellite (MIST), built at KTH Royal Institute of Technology. Two methods that make use of different techniques to measure the magnetic dipole moment of objects are evaluated through simulations. The first method holds the promise of being able to accurately estimate the dipole moment on components and the whole satellite alike, but has the downside of having a more complex setup. The second method can be set up easily, and can quickly produce an estimate of the dipole moment of one single object. However, the method is more susceptible to external disturbances in the magnetic field, and placement of the object. Due to time constraints, only the second method is evaluated experimentally. To understand how the second method performs, reference measurements are made on a coil with a known dipole moment. The results from the reference measurements show that the second method works well enough to produce values accurate enough for this project. Measurements are thereafter made on components similar to the flight hardware which are put through a set of tests to see how easily magnetised they are. The resulting values show that the magnetic field from magnetic tools can magnetise the components to the extent of becoming a problem, making the satellite’s dipole moment exceed the set limit. A more thorough investigation of MIST’s magnetic dipole moment should be conducted to determine if the satellites total magnetic dipole moment runs the risk of exceeding the set limit. / Små satelliter kan påverkas negativt av jordens magnetfält då det i samband med satellitens magnetiska dipolmoment kan resultera i ett moment som verkar på satelliten. Därför behöver det magnetiska dipolmomentet uppskattas under utvecklingen av satelliten för att fastställa att problem inte uppstår när satelliten är i omloppsbana runt jorden. Denna uppskattning måste göras på MIniture STudent satelliten MIST, som byggs på Kungliga Tekniska Högskolan i Stockholm. Två metoder som använder sig av olika tekniker för att mäta det magnetiska dipolmomentet på komponenter undersöks genom simuleringar. Första metoden verkar från simuleringarna kunna göra en mätning av det magnetiska dipolmomentet med hög precision på enskilda komponenter och hela satelliten. Metoden har dock en komplex mätuppställning. Andra metoden kan snabbt ställas upp och en uppskattning av en komponent kan fås. Nackdelen med metoden är dess känslighet mot störningar i magnetfältet samt felplacering av dipolen. På grund av projektets tidsram väljs den andra metoden för att göra experimentella mätningar. För att förstå hur den andra metoden presterar görs mätningar på en spole med ett känt magnetiskt dipolmoment. Resultaten från mätningarna på spolen visar att metoden fungerar bra nog för att ge tillräckligt noggranna värden för projektet. Mätningar görs därefter på komponenter som liknar MISTs hårdvara för att undersöka hur lätta de är att magnetisera. Resultaten visar att magnetiska fält från verktyg kan magnetisera komponenterna så pass mycket att den satta gränsen för satellitens magnetiska dipolmomentet överskrids. En mer grundlig utredning av MISTs magnetiska dipolmoment bör göras för att fastställa om dipolmomentet riskerar att överskrida gränsen. / Kandidatexjobb i elektroteknik 2020, KTH, Stockholm

A swept neutral pressure instrument(SNeuPI) is used to study the effect of gravity waves on the composition of the ionosphere. When mounted on a nanosatellite in the low earth orbit, changes in atmospheric pressure due to gravity waves are measured as the changes in neutral gas density. This measurement is achieved by use of micro-tip emitters as an electron source and micro channel plates(MCPs) as ion collectors. Ionization of the neutral gas produces a current at the output of the MCPs to quantify the pressure of the ionosphere. Traditionally, such measurements are made on larger satellites which enable the use of higher power equipment. This thesis describes the design and use of a low power instrument, to be used on a limited-resource satellite. The background and theoretical analysis is presented first, followed by descriptions of the mechanical and electrical designs. The laboratory tests are limited to a vacuum chamber setup that simulates the conditions of the ionosphere. / Master of Science